1. Field of the Invention
This invention relates to gas turbine engine rotor blades and, more particularly, to leading edges having localized compressive residual stresses imparted by laser shock peening.
2. Description of Related Art
Gas turbine engines and, in particular, aircraft gas turbine engines rotors operate at high rotational speeds that produce high tensile and vibratory stress fields within the blade and make the fan blades susceptible to foreign object damage (FOD). Vibrations may also be caused by vane wakes and inlet pressure distortions as well as other aerodynamic phenomena. This FOD causes nicks and tears and, hence, stress concentrations in leading and trailing edges of fan blade airfoils. These nicks and tears become the source of high stress concentrations or stress risers and severely limit the life of these blades due to High Cycle Fatigue (HCF) from vibratory stresses. FOD damage may also result in a loss of engine due to a release of a failed blade. It is also expensive to refurbish and/or replace fan blades and, therefore, any means to enhance the rotor capability and, in particular, to extend aircraft engine fan blade life is very desirable. The present solution to the problem of extending the life of fan blades is to design adequate margins by reducing stress levels to account for stress concentration margins on the airfoil edges. This is typically done by increasing thicknesses locally along the airfoil leading edge which adds unwanted weight to the fan blade and adversely affects its aerodynamic performance. Another method is to manage the dynamics of the blade by using blade dampers. Dampers are expensive and may not protect blades from very severe FOD. These designs are expensive and obviously reduce customer satisfaction.
Therefore, it is highly desirable to design and construct longer lasting fan blades that are better able to resist both low and high cycle fatigue than present fan blades. The present invention is directed towards this end and provides a fan blade with regions of deep compressive residual stresses imparted by laser shock peening leading and trailing edge surfaces of the fan blade.
The region of deep compressive residual stresses imparted by laser shock peening of the present invention is not to be confused with a surface layer zone of a work piece that contains locally bounded compressive residual stresses that are induced by a hardening operation using a laser beam to locally heat and, thereby, harden the work piece such as that which is disclosed in U.S. Pat. No. 5,235,838, entitled "Method and apparatus for truing or straightening out of true work pieces". The present invention uses multiple radiation pulses from high power pulsed lasers to produce shock waves on the surface of a work piece similar no methods disclosed in U.S. Pat. No. 3,850,698, entitled "Altering Material Properties"; U.S. Pat. No. 4,401,477, entitled "Laser shock processing"; and U.S. Pat. No. 5,131,957, entitled "Material Properties". Laser peening as understood in the art and as used herein means utilizing a laser beam from a laser beam source to produce a strong localized compressive force on a portion of a surface. Laser peening has been utilized to create a compressively stressed protection layer at the outer surface of a workpiece which is known to considerably increase the resistance of the workpiece to fatigue failure as disclosed in U.S. Pat. No. 4,937,421, entitled "Laser Peening System and Method". However, the prior art does not disclose fan blade leading edges of the type claimed by the present patent nor the methods how to produce them. It is to this end that the present invention is directed.